The invention relates to the internal cooling of gas turbine engine aerofoils and particularly but not exclusively to the cooling of turbine aerofoils.
Modern gas turbines operate with high turbine entry temperatures to achieve high thermal efficiencies. These temperatures are limited by the turbine vane and blade materials. Cooling of these components is needed to allow their operating temperatures to exceed the materials"" melting points without affecting the vane and blade integrity.
A large number of cooling systems are now applied to modern high temperature gas turbine vanes and blades. Cooling is achieved using relatively cool air bled from the upstream compressor system, the air bypassing the combustion chamber between the last compressor and first turbine. This air is introduced into the turbine vanes and blades where cooling is effected by a combination of internal convective cooling and external film cooling.
In film cooling a protective blanket of cooling air is ejected onto the external surface of the turbine vane or blade, from internal passages within the aerofoils, by means of holes or slots in the surface. The aim is to minimise the external heat transfer from the hot gas stream into the component surface.
In convective cooling the air is passed through passages within the aerofoil. This cools the metal since the air temperature is below that of the metal. Effectively the turbine component itself acts as a heat exchanger.
Unfortunately bleeding air from the compressor to cool the turbine reduces the overall cycle efficiency of the gas turbine engine. In addition, film cooling by ejecting air onto the turbine component surface causes aerodynamic losses in the turbine itself. Thus improvements in the performance of cooling systems continue to be soughtxe2x80x94either to cool the turbine at a given inlet temperature with less cooling air (improving cycle and turbine efficiencies), or to enable higher inlet temperatures to be sustained with the existing levels of cooling air consumption.
According to the invention, there is provided an aerofoil for a gas turbine engine, the aerofoil including an elongate internal cooling passage for receiving a flow of cooling fluid and an elongate internal feed passage extending at least partially alongside the cooling passage, the cooling passage and the feed passage being separated by an elongate internal wall, wherein a plurality of openings are provided in the wall for feeding cooling fluid from the feed passage into the cooling passage, to induce at least two vortices in cooling fluid flowing through the cooling passage.
Preferably the openings are angled such that fluid flowing therethrough has a component of movement in a direction parallel to the cooling passage.
Preferably, the internal wall includes two sets of openings, each set including a plurality of openings generally aligned in a direction parallel to the cooling passage, and each set of openings providing means for inducing a vortical flow of fluid in the cooling passage.
Preferably each set of openings extends along substantially the whole of the length of the cooling passage. The openings may be positioned so as to induce two generally parallel, adjacent vortices.
The cooling passage may be bounded along its length by further elongate walls, the further walls having substantially no openings therein, and at least one wall comprising a part of an outer wall of the aerofoil. Preferably the openings are located and oriented such that fluid flowing into the cooling passage initially flows along an inner surface of the outer wall of the aerofoil. Preferably one set of openings is oriented and located such that fluid flowing therethrough and into the cooling passage initially flows along the inner surface of a wall forming a suction side wall of the aerofoil and the other set of openings is oriented and located such that fluid flowing therethrough and into the cooling passage initially flows along the inner surface of a wall forming a pressure side wall of the aerofoil.
One set of openings may be located and oriented to induce a vortex which rotates in a first direction and the other set of openings may be located and oriented to induce a vortex which rotates in the opposite direction.
Preferably fluid within one vortex flows initially along the inner surface of the wall forming a suction side wall of the aerofoil and subsequently along an internal wall of the aerofoil and fluid within the other vortex flows initially along the inner surface of the wall forming a pressure side wall of the aerofoil and subsequently along the same internal wall of the aerofoil, the two fluid-flows meeting at a central region of the internal wall.
The openings in the wall may be located and oriented to induce a vortex having a screw-type motion, with a component of movement in a direction parallel to the cooling passage. Inner surfaces of walls of the cooling passage may be provided with ribs aligned with the screw-type path of motion of the fluid within the vortex.
Preferably, the feed passage is located in a leading or trailing edge of the aerofoil and the cooling passage is located in an internal region of the aerofoil. The aerofoil may include a feed passage at its leading edge, a feed passage at its trailing edge and two cooling passages located therebetween, each cooling passage being fed with cooling fluid from an adjacent feed passage.
Preferably the aerofoil is adapted to be oriented in a generally radial direction of the gas turbine engine and the cooling passage extends generally in the radial direction of the gas turbine engine when the aerofoil is so oriented. The aerofoil may comprise a part of a turbine blade for the gas turbine engine, adapted to be mounted on a rotor disc so as to extend radially therefrom. The turbine blade may include a root portion for mounting on the disc, the root portion including a passage through which fluid may pass to the feed passage.
Alternatively the aerofoil may comprise a part of a turbine stator or a nozzle guide vane for the gas turbine engine.
According to the invention, there is further provided a gas turbine engine including an aerofoil according to any of the preceding definitions.
According to the invention, there is further provided a method of cooling an aerofoil for a gas turbine engine, the aerofoil including an elongate internal cooling passage, wherein the method includes the step of providing a flow of cooling fluid in the passage and inducing at least two vortices in the fluid. The fluid within each vortex may have a screw-type motion, with a component of movement in a direction parallel to the cooling passage.